XFOIL Version 6.94 Calculated polar for: pp14 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0033 0.04675 0.03812 -0.0161 0.9996 1.0004 -2.750 -0.0041 0.04579 0.03731 -0.0143 0.9996 1.0004 -2.500 -0.0053 0.04483 0.03650 -0.0125 0.9996 1.0004 -2.250 -0.0070 0.04388 0.03569 -0.0106 0.9996 1.0004 -2.000 -0.0090 0.04292 0.03487 -0.0086 0.9996 1.0004 -1.750 -0.0112 0.04194 0.03405 -0.0067 0.9996 1.0004 -1.500 -0.0260 0.04101 0.03328 -0.0021 0.9996 0.9941 -1.250 -0.0611 0.04016 0.03255 0.0067 0.9996 0.9785 -1.000 0.2590 0.05147 0.04318 -0.0829 0.9996 0.6419 -0.750 0.3911 0.05445 0.04570 -0.1146 0.9996 0.4769 -0.500 0.4485 0.05462 0.04577 -0.1240 0.9996 0.4254 -0.250 0.5017 0.05513 0.04612 -0.1319 0.9996 0.3891 0.000 0.5453 0.05538 0.04629 -0.1369 0.9996 0.3649 0.250 0.5864 0.05583 0.04664 -0.1411 0.9996 0.3453 0.500 0.6235 0.05625 0.04704 -0.1442 0.9996 0.3328 0.750 0.6581 0.05689 0.04768 -0.1464 0.9996 0.3259 1.000 0.6904 0.05747 0.04838 -0.1479 0.9996 0.3263 1.500 0.9918 0.04236 0.02967 -0.1632 0.2531 0.4125 1.750 1.0164 0.04325 0.03076 -0.1617 0.2411 0.4669 2.000 1.0379 0.04353 0.03153 -0.1593 0.2337 1.0004 2.250 1.0756 0.04508 0.03198 -0.1587 0.2276 1.0004 2.500 1.1168 0.04665 0.03312 -0.1592 0.2222 1.0004 2.750 1.1680 0.04856 0.03473 -0.1615 0.2167 1.0004 3.000 1.2309 0.05126 0.03715 -0.1662 0.2107 1.0004 3.250 1.2791 0.05353 0.03968 -0.1684 0.2084 1.0004 3.500 1.3238 0.05632 0.04271 -0.1701 0.2091 1.0004 3.750 1.3626 0.05959 0.04622 -0.1711 0.2108 1.0004 4.000 1.3900 0.06211 0.04934 -0.1700 0.2140 1.0004 4.250 1.4119 0.06522 0.05318 -0.1685 0.2192 1.0004 4.500 1.4339 0.06911 0.05756 -0.1674 0.2247 1.0004 4.750 1.4556 0.07353 0.06228 -0.1666 0.2298 1.0004 5.000 1.4698 0.07756 0.06683 -0.1649 0.2351 1.0004